Superconductive Hypersonic Liquefaction Nosecone

ABSTRACT

An apparatus and method for mitigating the shock front of a rocket or aerospace plane flying at hypersonic speeds while simultaneously distilling liquid chemical elements from the ambient air. The invention employs supercooling driven by the cryogenic power of liquid hydrogen and/or regenerative evaporation of liquid hydrogen and/or liquid nitrogen to drive isothermal compression and consequentially usurp the shock front in totality.

CROSS REFERENCE TO RELATED APPLICATIONS

The present application is a Divisional of U.S. Utility patentapplication Ser. No. 13/969,467 filed on Aug. 16, 2013. Which in turnclaims the benefit of U.S. Provisional Patent Application No.61/683,836, filed on Aug. 16, 2012 and entitled “SuperconductiveHypersonic Liquefaction Nosecone,” which is herein incorporated byreference in its entirety.

BACKGROUND

This invention relates to the nosecone design of a rocket or aerospaceplane. More specifically, this invention relates to the mitigation ofthe shock front experienced at hypersonic speed through supercooling.

U.S. Pat. No. 5,191,761, owned by the applicant for the presentinvention, discloses an air breathing aerospace engine. That patent isincorporated by reference in its entirety. The engine includes a frontalcore that houses an oxygen liquefaction system that captures ambientair, liquefies and separates the oxygen. The oxygen may then be used inthe rocket engine.

U.S. Pat. No. 6,213,431, owned by the applicant of the presentinvention, discloses an aerospike engine. That patent is incorporated byreference in its entirety. An aerospike engine may have a tapered bodywith a slanted or curved reaction plane. A fuel injector directs fueldown the reaction plane. The combustion of the fuel on the reactionplane creates a propulsive force across the reaction plane.

U.S. Pat. No. 7,344,111 hence discloses a re-usable or reversible SSTOthat may be expediently launched to service the rapidly expanding spaceenterprise. Merging patent '761 and patent '431 rendered a uniqueaerospace plane confirmation (the '111 patent) with a liquefaction orsupercooling nosecone on the ascending or leading direction and anadaptive aerospike rocket engine on the tail end.

Oxygen constitutes 89% of the mass of the necessary propellant for suchan engine. A substantial portion of the necessary liquid oxygen can bedistilled out of the ambient atmosphere at hypersonic speeds bysupercooling through turbo compression at the nosecone. Through thisprocess, the merged aerospace plane would be capable of tanking thenecessary hydrogen as well as carrying a substantially improved payloadinto orbit. However, a consequence of tanking all of the necessaryliquid hydrogen propellant onboard would be to invariably result inbubbly contours.

A reversible (“Uturn”) aerospace plane was rendered that can flyhypersonically without the impediment of shock waves, shock front, orsuperheated intake air (a node of optimality) in the ascending directionwhilst reentering in the reversed (high drag) direction into theatmosphere with the benefit of cushioned heat dissipation in thehigher/rarified atmospheric domain. The slated “turbocharged” aerospaceplane may also be barrel-rolled through the initial reentry phase so asto dissipate the insipient heat into deep space. As a consequence ofemploying the cryogenic potential of the liquid oxygen/hydrogenpropellants to supercool the nosecone in lieu of cooling the jacket ofthe rocket engine, it would be conversely be necessary to line theexpansion ramp of the aerospike engine with ceramic tiles. Slatedceramic tiled would function both as a reentry shield as well as activeramp insulation. The Uturn would hence facilitate an aerospace planewith substantially lesser heat shield whilst limiting ceramic tiles tothe aerospike engine expansion ramps.

Whereas the efficacy of the oxygen liquefaction would be substantiallyconstrained by the extremely low pressure of the rarified atmosphericair at high altitudes with the “111” patent in its native form, theaddition of a turbo compressor driven by either superheated nitrogen orsuperheated hydrogen in accordance with claims #1 and #13 of the “111”patent would substantially expand the operating envelope of the airbreathing aerospace engine. The high compressed ambient air may hence beinter-cooled in accordance with the means provided via patent “111” andhence expanded/flashed into the cryogenic zone via an expansionmotor/turbine that may either augment the compression process or drivethe propellant pumps. Means of supercooling the nose cone at supersonicspeeds is additionally being introduced to suppress the formation of theshock front in the abstract (eg lim dQ>>infinity, dL=0), “freezing” ORusurping the incipient shock at Mach 1 and rendering isothermalcompression in lieu of adiabatic compression of the intake air into theliquefaction plant throughout the hypersonic regime.

A novel new approach with a rocket or aerospace plane flyinghypersonically through the atmosphere entails morphing (cool) isothermalcompression (in lieu of a blazing shock front) by orchestratingsupercooling and superconductivity in coherence, usurping the insipientshock wave at formation (Prandtl singularity, Mach1) and distilling bothliquid oxygen as well as liquid nitrogen simultaneously out of theincident atmospheric air via regenerative chilling. The rational simplyentails that in lieu of dissipating 99.99% of the incident kineticenergy in the shock front at hypersonic velocity (an irreversibleadiabatic process), 99.99% of the kinetic energy is conversely convertedinto useful work via isothermal compression of the incident atmosphericair. By employing the cryogenic power of tanked hydrogen incident to aH2/O2 propulsive system coherently and morphing isothermal compressionby means of the Prandtl singularity “niche” via the force ofsupercooling AND superconductivity, compressedsupercool/saturated/liquefacted ambient air may be rendered as a uniquesolution to a perplexing technological challenge.

The following definitions apply to terms used in this application:

-   -   Liquefaction: The condensation of a gaseous medium.    -   Supercooling: Rate of cooling orders of magnitude in excess of        normal heat transfer rates.    -   Superconduction: Rate of conduction orders of magnitude in        excess of normal conduction rates.    -   Superemissivity/absorptivity: Emissivity or absorptivity in        excess of unity by means of negative refraction (or contact area        morphing).    -   Supersonic: Flying 2-3 times the speed of sound.    -   Hypersonic: Flying 4-20 times the speed of sound.    -   Isothermal compression: Compression at constant temperature    -   Adiabatic compression: Compression without loss of heat (eg        rapid compression via a shock wave).

Normally compression of air results in an increase in the temperature ofthe incident air. Air may be compressed either via a mechanical device(reciprocating or turbo compressor) OR via the ram force of a bodytraveling though the atmosphere. Compression of the air may be enhancedvia a diffuser in the latter instance. At high speeds rapid compressionof the air results in formation of a shock wave (adiabatic/trapped heatof compression). The speed of the body traveling through the atmosphereat formation of a shock wave is denoted Mach1. The energy content of theair at Mach1 is labeled “total” or “stagnation” temperature “Ts” inabsolute terms. The pertinent relationships are as follows:

Ts=To[1+(k−1)/2×M̂2]  (1)

where Ts is the stagnation temperature, To=ambient temperature inabsolute terms, k=polytropic constant Cp/Cv=1.4 and M=Mach number;

T2/T1=(p2/p1)̂k−1/k  (2)

Where T2/T1 is the adiabatic temperature ratio and p2/p1 is thecompression ratio;

wa=Cp·R·T1·(k/k−1)[(p2/p1)̂k−1/k−1]  (3)

where wa is the adiabatic work of compression, R is the ideal gasconstant, T entering temperature;

wi=Cp·R·T1·ln(p2/p1)  (4)

where wi is the isothermal work of compression with a compression ratioof p2/p1 vis-à-vis; Atmospheric pressure model(http://en.wikipedia.org/wiki/Atmospheric_pressure);

1 0 ½  18,000 ft ⅓  27,480 ft 1/10  52,926 ft 1/100 101,381 ft 1/1,000159,013 ft 1/10,000 227,899 ft 1/100,000 283,076 ft. 0/E{circumflex over( )}3 = 20; 50,000 ft/E{circumflex over ( )}6 = 403; 95,000ft/E{circumflex over ( )}9 = 8,103; 150,000 ft/E{circumflex over ( )}12= 162,755The real time Mach number/stagnation temperature relationship in termsof altitude has been determined as follows(http://www.grc.nasa.gov/WWW/BGH/stagtmp.html);

-   -   M1=500 R @5000 ft    -   M2=700 R @30-90,000 ft    -   M3=1100 R @40-90,000 ft    -   M4=1600 R @40-90,000 ft    -   M5=2200 R @40-100,000 ft    -   M6=3200 R @40-100,000 ft    -   M7=4200 R @40-100,000 ft    -   M8=5300 R @40-100,000 ft.        The respective shock/gradient for a (real time) adiabatic shock        front (plus associated frontal pressure) may hence be rendered        via the polytropic relationship p2/p1=[T2/T1]̂k/(k−1) as follows;

M1: p2/p1 = 2.2 p2 = paxp2/p1 = 2 × 15/2.75 = 11 psi @20,000 ft M2:p2/p1 = 3.2 p2 = paxp2/p1 = 3.2 × 15/4 = 12 psi @30,000 ft M3: p2/p1 =16 M4: p2/p1 = 59 M5: p2/p1 = 179 p2 = paxp2/p1 = 179 × 15/55 = 49 psi@70,000 ft M6: p2/p1 = 663 M7: p2/p1 = 1718 M8: p2/p1 = 3878 ×15/1100 =53 psi @100,000 ft In event of isothermal compression the Bernoulli's law will control withp=V̂2/2vg/144 in psi. The respective frontal pressure in event ofisothermal compression hence becomes:

-   -   M1: p1=    -   M2: p2=10.3 psi @27,480 ft (½)    -   M3: p3=    -   M4: p4=3.9 psi @52,926 ft (⅓)    -   M5: p5=    -   M6: p6=2.8 psi @101,381 ft ( 1/100)    -   M7: p7=    -   M8: p8=2.0 psi @ 120,000 ft ( 1/245)    -   M9: p9=    -   M10: p10=0.77 psi @ 159,013 ft ( 1/1000)        The work of compression for a (normalized) free-range adiabatic        compression hypersonic shock ram versus isothermal compression        (in the abstract) at Mach8 is as follows:

wa = Cp.R.T 1.(k/k − 1)[(p 2/p 1)⋀k − 1/k − 1] = 0.25 × 3.5 × 53.3 × 400/788 × [(3878)⋀0.286 − 1] = 226  Btu/lbwi=Cp·R·T1·ln(p2/p1)=0.25×53.3×400/788×ln(245)=37 Btu/lb.

The impact of a sustainable isothermal (ram) compression system thru thehypersonic domain is threefold, eg 1) that the destructive thermalimpact in the native/adiabatic mode is being contained 2) the work ofcompression is substantially mitigated to the extent that the latentheat of evaporation of hydrogen component alone of the LH2/LO2propellant will suffice in dissipating the incident heat of compressionwithin the realm of the supercool synthesis and 3) the drag force of a“superconductive hypersonic liquefaction nosecone” will be proximal 6×(eg 226/37) less in comparison to a native hypersonic aerospace planeflying trough the hypersonic zone.

The mass flow rate at Mach8 (8,000 ft/sec)conversely=8,000×1/14/245=2.33/lb/SF/sec. The heat of compression hencebecomes:

-   -   Qa=226×2.33×3600=1,895,688 Btu/SF/h    -   Qi=37×2.33×3600=310,356 Btu/SF/h.    -   O2: Latent Heat of Vaporization: 2934 BTU/lb mole (92 Btu/lb)        http://www-safety.deas.harvard.edu/services/oxygen.html#physical    -   N2: Latent Heat of Vaporization: 2399 BTU/lb mole (86 Btu/lb)        http://www-safety.deas.harvard.edu/services/nitrogen.html#physical    -   H2: Latent Heat of Vaporization: 389 Btu/lb-Mole (195 Btu/lb)        http://www.ehs.ufl.edu/Lab/Cryogens/hydrogen.html

H2 has two different nuclear spin states: ortho-hydrogen (spin 1 state)and para-hydrogen (spin 0 state). Near room temperature, equilibriumhydrogen is 75% ortho and 25% para. However, at low temperatures, aroundthe normal boiling point of 20.3 K, hydrogen is nearly all in thepara-hydrogen state. The conversion process from ortho to para hydrogenis exothermic and generates about 700 KJoule/mole.

SUMMARY

The invention comprises means to supercool the incipient shock wave, ofa rocket or (aero) space plane. Supercooling (or superchilling) willresult in usurpation of the shock wave by dissipating the heat ofcompression at the same rate of formation. By dissipating the heat ofcompression at an infinite rate throughout the supersonic and hypersonicregime, the shock front and associated heat would be totally abated.During dissipation of the shock front, the ambient air at inception willbe isothermally compressed throughout the supersonic and hypersonicregimes. Cool isothermal compression (as opposed to hot adiabaticcompression) is an essential prerequisite for efficient liquefaction ofambient air. Liquid hydrogen may be employed to chill the nosecone of arocket or space plane to mitigate the shock front at hypersonic speed.Supercooling, on the other hand, through a combination ofsuperconductivity and superemissivity, may usurp the shock front in itstotality at Mach 1. The supercooling is driven by the cryogenic power ofliquid hydrogen, regenerative flash evaporation of liquid hydrogenand/or liquid nitrogen, and consequential isothermal compression. Theresulting supercooled air may be compressed, regeneratively intercooled,and flashed into liquid air. Under conditions of optimality theincipient air stream may be compressed in accord with the stagnationpressure of the Mach number and liquefacted in one-step via the power ofsupercooling and isothermal compression.

In order to achieve supercooling, a superconductive heat exchangermedium will be required. Although copper becomes a superconductive atthe flashing temperature of liquid hydrogen, copper in isolation is nota suitable material for a superconductive heat exchanger driven byliquid hydrogen. Liquid hydrogen would also be a hazardous supercoolingmedium, as leakage would result in destruction. This invention providesfor a novel approach employing liquid nitrogen (an inert medium) and adouble-shelled, glass dome with superemissive coatings on the inside andoutside to facilitate supercooling through superemissive/superabsorptivecoatings. Superconductivity will hence be morphed through (1) 4^(th)power driving force of the Boltzman equation system and (2) closecoupling of the radiative surfaces. The glass surfaces may simply beetched or coated with special means (for example diamond power) so as tofacilitate the desired emissivity/conductivity.

In one aspect the invention relates to a rocket or aerospace planeflying hypersonically through the atmosphere with means tomitigate/eliminate the shock front at hypersonic speed via the cryogenicpower of liquid hydrogen. In another aspect the invention relates to amethod of distilling oxygen out of atmospheric air by means of thecooling capacity of liquid/slush hydrogen. In a third aspect of theinvention, the Prandtl singularity is introduced as the route into theabstract (for example changing lanes) to mitigate/eliminate the shockfront at hypersonic speed. A fourth component of the invention ismorphing superconductivity as superemissivity through a concentric,close-coupled (double-decker) black bulb radiation system driven byliquefacted nitrogen. A fifth aspect of the invention constitutesmorphing of superemissivity by superemissive form factor through“scatter” emissivity.

A sixth aspect of the invention introduces liquid hydrogen as theprinciple cryogenic potential force. A seventh aspect of the inventioncomprises the power of supercooling and superconductivity in sink asmeans to open the door of the Prandtl singularity. An eighth aspect ofthe invention comprises the cooling/chilling force of the liquefactedoxygen and liquefacted nitrogen. A ninth aspect of the inventioncomprises maximizing the supercooling and liquefaction potential of boththe liquefacted oxygen and liquefacted nitrogen through regenerativecooling.

Whereas liquid hydrogen on its own merits may be rationally applied tomitigate the shock front at hypersonic speed and distilled liquid oxygenin a one-step process, the efficacy is limited to a narrow band of thehypersonic domain. However, lining up the powers of supercooling,superconductivity, super emissivity and regenerative cooling opens thedoor to the abstract: a rocket or aerospace plane flying hypersonicallythrough the atmosphere without impediment through the use of asuperconductive hypersonic liquefaction nosecone.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates an aerospace plane with an aerospike rocket engine onthe tail end and a conical/spiked nosecone incorporating a louveredintake aperture with concentric rings on the leading end as an extensionof the solid cone.

FIG. 2 illustrates an aerospace plane with an aerospike rocket engine onthe tail end of a spherical nosecone incorporating a radial louveredintake aperture comprising concentric rings on the leading end.

FIG. 3 illustrates an enlargement of the spherical nosecone(supercool/superconductive probe) and radial (incident air) intakelouvers (regenerative liquefaction heat exchanger). It also illustratesshedding of ice emanating from ambient condensate.

FIG. 4 illustrates the makeup of the double-decker, super-emissive (or“scatter”) blackbulb spherical nosecone in concert with the regenerativeradial intake louvers. It also illustrates the relevance of liquid andliquefacted nitrogen as supercooling medium.

FIG. 5 illustrates the active components, like the optically black innersphere or shell, and external etching or coating of the outer glasssphere or shell of the double-decker, super-emissive blackbulb sphericalnosecone.

FIG. 6 illustrates an alternative finned tube (black bulb) approach as asubstitute to the black bulb anchor with a solid inner shell.

DETAILED DESCRIPTION

A rocket or aerospace plane in accordance with one or more embodimentsof the invention construes an aerospace plane flying hypersonicallythrough the atmosphere with means to mitigate/eliminate the shock frontat hypersonic speed via the cryogenic power of liquid hydrogen or liquidmethane via the (Prandtl singularity) at Mach 1. Liquefacted oxygen mayhence be harvested as a consequence of rational alignment of thesupercooling potential of (tanked) liquid hydrogen and regenerativepower of liquefacted oxygen and liquefacted nitrogen. As an extension ofthe premise of liquefaction/regeneration, liquefacted nitrogen may beapplied in substitution of (tanked) hydrogen and/or liquid methane asthe supercooling denominator in cohesion with structuredsuperemissivity.

More specifically in accordance with the premise of the invention of arocket or aerospace plane flying hypersonically through the atmosphere,FIG. 1 illustrates a simplistic aerospace plane 100 with an aerospikerocket engine 180 on the tail end and a conical/spiked nosecone 101 onthe leading end. The nosecone incorporates a louvered intake aperture102 with concentric rings on the (leading) end as an extension of thesolid cone entraining incident/atmospheric air 105. FIG. 1 alsoillustrates adaptive aerospike keys 181 and 182, the aerospace planebody 190, a parafoil wing 191, a leading edge 192 and shroud/cowling103. Cowling 103 may be construed as a cooling shroud as well as asalient regenerative nitrogen (ejector) heat exchanger 106. Functionallythe incident ambient air 105 flows over nosecone 101 and entersspacecraft 100 via aperture 102. Normally nosecone 101 would act as anablative shock cone. However with the advent of the art of supercoolingand superconductivity, nosecone 101 will be morphed into asuperconductive hypersonic liquefaction nosecone at hypersonic speed,usurping the incident shock front at formation whilst rendering (as abyproduct) a highly compressed/saturated air stream that will enterspacecraft 100 via the louvered aperture 102.

FIG. 2 conversely illustrates another embodiment of the invention 100whereby the nosecone 101 is configured as a spherical probe 110 and thelouvered aperture 102 (from FIG. 1) as a finned cylinder 120.Functionally the incident ambient air 105 flows over spherical nosecone110 (supra nosecone 101, FIG. 1) and enters spacecraft 100 viacylindrical aperture 120 (supra aperture 102, FIG. 1). Because of theextended surface area nosecone 110 will be superior to nosecone 101 dueto enhanced ability to usurp the incident shock front at formation.Nosecone 110 also has a greater surface area and extendedcircumferential exposure with regards to the conical format of nosecone101. More specifically the extended surface area of spherical nosecone110 is an essential denominator with regards to the premise ofoptimality, eg limit dQ>>infinity, dL=0. Whereas dL constitutes theboundary layer, it is imperative to max the circumferential area tosatisfy the condition of optimality to open the door to the Prandtlsingularity, usurping the incident shock front and renderingcool/saturated (isothermally) compressed fluid stream entering aperture120.

FIG. 3 illustrates the parametric functioning of spherical nosecone 101more specifically. More specifically the incident air is substantiallycompressed into a minute boundary layer. Compression of the boundarylayer 108 is an imperative prerequisite in accordance with the premiseof optimality (eg dL=0). Hence conforming cone 101 (FIG. 1) into sphere110 (FIGS. 2/3) in maximizing the circumferential area. The same appliesto maximizing the length of the fluid path (area of contact) that theincident air 105 must follow around sphere 110 prior to enteringaperture 120. Max contact is a prerequisite for optimality. Sphere 110offers maximum area of contact. Sphere 110 may however be dimpled,fluted, warped or finned to extend the contact area and to offer adrainage route for liquefacted oxygen. FIG. 3 also illustrates liquidhydrogen as the primary supercooling agent. Sphere 110 is converselyconstructed out of copper that becomes superconductive at −420 F/40 R.Because of the superconductivity of sphere 110 and supercool expansionof liquid hydrogen (−423 F/37 R) the rate of heat transfer dQ would beinfinitely high (a prerequisite for the condition of optimality ofdQ>>infinity). Shedding of ice 108 would also be enhanced by thetangential inversion of the direction of flow of incident air 105 aroundsphere 110.

FIG. 4 illustrates the art of morphing copper sphere 110 (from FIG. 3)into a double-decker glass sphere 112 and (inner-concentric) coppersphere 114. Pointer 106 illustrates the close-contact prerequisite.Rationale for concentric spheres is three-fold, eg 1)isolating/insulating the supercool copper sphere 114 from the ambientenvironment 2) creating the ability to regulate the outer (glass) sphere112 at an intermediate temperature and 3) to substitute liquefacted(regenerated) nitrogen as the principal supercooling medium by employingby means of a (finned) shell and tube inner sphere 114 in lieu ofcopper. The rationale (A) is simply (1) by virtue of an optically blacketched/coated outer sphere 112 (2) maintaining the inner sphere 114optically black and (3) radiation heat transfer driven by the 4th powerof the absolute temperature, the rate of heat transfer would beinfinitely high in event of an 100 C of 180 F temperature differentialbetween the outer and inner spheres. The glass sphere would also becomea passive component in the transfer process with externaletching/coating of the external sphere. Rationale (B) is much morecompelling inasmuch as substituting the inner (copper) sphere with aspherical finned coil bundle, liquefacted nitrogen may be employed asthe principal supercooling medium as an optical black fin-coil bundlewould for all practical purposes constitute a perfect black bodyobviating the need for a superconductive copper shell.

FIG. 5 illustrates the “double-decker” (black bulb) synthesis morepunctually via superemissive concentric spheres. More specifically heatis radiated from the (optically black) etching/coating 131 of the outersphere 130 (through the glass shell) to the (optically black)etching/coating of the inner sphere 134. Superemissivity is driven via(imbedded) micro/nano prisms 136. The spheres are in close proximity toeach other.

Radiation heat transfer is governed by the Boltzmann equation:

QR=A×AZ×EZ×BZ×(T2̂4−T1̂4)

where A=area, AZ=form factor, EZ=emissivity and BZ=Boltzmannconstant=0.1714×10-8 Btu/hour/ft2/R4.

Normally both the form-factor (radiative contact) and the emissivity(optical contact) range from 0-1.0. In event of a tight fit=1.0. Sametoken EZ=1.0 for black bodies. In event T2=460 R (source) and T1=140 R(sink) the rate of heat transfer QR becomes QR=1×1×1×0.1714×10̂-8×460̂4(Btu/SF/h)=0.1714×(10̂-8)×447.8×10̂8=77.9 Btu/SF/h (in the limit T1̂4=0).At 600 R (source) the rate of heat transfer QR=222.1 Btu/SF/h, at 800 R(source) QR=702 Btu/SF/h and 1000 R (source) (560 F) QR=1,714 Btu/SF/h.Since a dissipation rate of 310,356 Btu/SF/h is required so as tomaintain isothermal compression in terms of the Prandtl singularity, the“superemissive” double-decker (superconductive) model will fail on theface. The solution to the superemissive-superconductivity quest hencevests with morphing of a “superemissive” form-factor. The solution isfound in “scatter” emissivity via the etching or coating of micro/micronprisms imbedded on both the (active) source and sink surfaces. Given forexample 10 micron prisms with a 45 deg cut-off and 50% spectral efficacyand 50% translucency, a multiplication factor of(10̂6/10)×0.5×0.5×0.5=1,563 is being rendered. The respective“SUPEREMISSIVE” transfer rates hence become:

-   -   460 R: QRR=77.9×1563=121758 Btu/SF/h    -   600 R: QRR=222×1563=346,986 Btu/SF/h    -   800 R: QRR=702×1563=1,097,226 Btu/SF/h    -   1000 R: QRR=1714×1563=2,678,982 Btu/SF/h.

Mastering the art of scatter emissivity hence constitutes the key tomorphing supercooling via constructive superemissivity. It is the key toopening the door to the Prandtl singularity and controlling thehypersphere rationally via isothermal compression and (regenerative)expansion of liquefacted nitrogen (and oxygen to a lesser extent) asprincipal supercooling agent(s).

FIG. 6 illustrates the art of morphing superconductivity via asuperemissive concentric finned radiator 135 (in lieu of concentricspheres) driven by micro/nano prisms 136 in concert with sphere 130 andetching/coating 131 in accordance with the double-decker (black bulb)scatter synthesis. With the radiator approach superemissivity may bemorphed because of the finned cavities may act as a perfect black bulbwithout special coatings. The liquefacted nitrogen and/or oxygen may beducted in a tube array in lieu of filling a sphere in totality.

Persons skilled in the art will recognize that many modifications andvariations are possible in the details, materials, and arrangements ofthe parts and actions which have been described and illustrated in orderto explain the nature of this inventive concept and that suchmodifications and variations do not depart from the spirit and scope ofthe teachings and claims contained therein.

All patent and non-patent literature cited herein is hereby incorporatedby references in its entirety for all purposes.

While the applicant understands that claims are not a necessarycomponent of a provisional patent application and has not includeddetailed claims, the inventor reserves the right to claim, withoutlimitation, the following subject matter.

I claim:
 1. An aerospace plane flying hypersonically through theatmosphere comprising: a body; a nosecone wherein are housed means tosimultaneously usurp the shock front and distill a liquid chemicalelement from the ambient air; and a rocket engine on a tail end.
 2. Theaerospace plane of claim 1 wherein the chemical element distilled fromthe ambient air is oxygen, nitrogen, or both.
 3. The aerospace plane ofclaim 1 wherein said means are accomplished through a superconductivityprocess.
 4. The aerospace plane of claim 3 wherein saidsuperconductivity process comprises of: dissipating a heat ofcompression of an incipient shock wave at formation at an infinite rate,isothermally compressing an incident air; and causing liquefaction of aportion of the incident air.
 5. The aerospace plane of claim 1 whereinsaid means are accomplished through a superemissivity process.
 6. Theaerospace plane of claim 5 wherein said superemissivity processcomprises of: dissipating the heat of compression of the incipient shockwave at formation at an infinite rate, isothermally compressing theincident air; and causing liquefaction of a portion of the incident air.7. The aerospace plane of claim 3 wherein the superconductivity processis driven by the expansion of tanked liquid hydrogen.
 8. The aerospaceplane of claim 3 wherein the superconductivity process is driven by theexpansion of liquefacted oxygen.
 9. The aerospace plane of claim 3wherein the superconductivity process is driven by the expansion ofliquid methane.
 10. The aerospace plane of claim 3 wherein thesuperconductivity process is driven by the expansion of liquefactednitrogen.
 11. The aerospace plane of claim 5 wherein the superemissivityprocess is driven by the expansion of tanked liquid hydrogen.
 12. Theaerospace plane of claim 5 wherein the superemissivity process is drivenby the expansion of liquefacted oxygen.
 13. The aerospace plane of claim5 wherein the superemissivity process is driven by the expansion ofliquid methane.
 14. The aerospace plane of claim 5 wherein thesuperemissivity process is driven by the expansion of liquefactednitrogen.